1. Field of the Invention
The present invention generally relates to deployable structures such as solar arrays and, more particularly, to solar array support systems.
2. Description of the Related Art
In outer space applications, solar arrays are generally composed of a series of solar panels for generating electrical power for systems such as a spacecraft or the like. The conversion of solar energy into electrical energy through solar cells is an obvious choice for producing power for such systems. The solar arrays are typically designed in foldable configurations in which the solar panels, supporting solar cells, are hingeably connected edge to edge lengthwise by various attachment systems. Such solar arrays may also include reflectors to concentrate the solar light upon the solar arrays. Reflectors are attached to the solar panels widthwise to the opposite ends of each solar panel.
The solar arrays are generally mounted on deployment yokes so that they can be extended or retracted from the spacecraft. During the launch of the spacecraft, the solar arrays are put into a stowed configuration where the solar arrays are folded in zigzag fashion against the spacecraft. Once the spacecraft is in outer space, the solar arrays are deployed into an extended configuration where the solar panels and the reflectors are folded away into an operation position in which the solar cells face the sun.
Such solar arrays must be adequately designed to withstand the undesirable physical conditions of such space missions so that they can properly function throughout their life time. Such undesirable conditions are generally mechanical and thermal stresses occurring during the launching and during the operation of the solar arrays. In this respect, the solar panels supporting the solar cells must be designed to meet the certain thermal and mechanical stress and strain requirements so as to protect the solar cells on them. As the solar cells are made of silicon or gallium arsenide materials, they are brittle. In other words, the panels function as the mechanical and the thermal support of the solar arrays.
Currently, solar panels are constructed from aluminum honeycomb substrates. In such structures, the honeycomb substrates are covered with carbon fiber face sheets on upper and lower surfaces of the honeycomb substrates. The carbon fiber face sheets stiffen the honeycomb substrate to increase the strength and the rigidity of the solar panels in stowed or deployed configurations. However, such aluminum honeycomb base solar panels are heavy in the context of such space applications. Another drawback involves their deployment and stowing systems. Such solar arrays require complex deployment and stowing systems employing tension wires, springs, hinges and the like to facilitate the deployment and stowing of the solar arrays.
As can be seen, there is a need for light weight, high temperature resistant, stiff and resilient deployable structures such as solar arrays, reflectors, and thermal blankets.
The present invention provides a deployable structure, such as a solar array system, utilizing a carbon fiber mesh material as a substrate. Due to its light-weight and flexibility, substrates containing the carbon fiber mesh material can be applied, as an example, to deployable solar arrays and reflectors for air borne vehicles such as satellites for space based applications as well as to the solar arrays for use in stratospheric platforms of an air borne vehicle such as an airplane.
In one aspect of the present invention, a solar array comprises a mesh substrate and a plurality of solar cells disposed on a first surface of the mesh substrate. The mesh substrate is formed from a matrix of resilient fibers.
These and other features, aspects and advantages of the present invention will become better understood with reference to the following drawings, description and claims.